1. Field of the Invention
The invention relates to stability augmentation systems particularly with respect to requirements for relaxed static stability aircraft.
2. Description of the Prior Art
As is known, conventional modern day aircraft are structurally designed to provide inherent aerodynamic longitudinal static stability so as to impart safe and desirable flight and handling characteristic to the aircraft. Typically this may be achieved by balancing the pitching moment of the center of lift about the center of gravity of the aircraft with an equal and opposite pitching moment from the horizontal stabilizer maintained in a deflected attitude with respect to its streamlined position. When, for example, a gust disturbance deflects the aircraft in pitch, an aerodynamically statically stable craft tends to return to its original attitude. When the pilot maneuvers such an aircraft in pitch by utilizing the control column, the aircraft responds to the command, holds a new angle of attack (which is approximately equal to an attitude change for most maneuvers) as long as the control column is deflected, and returns to its original angle of attack when the control column is released. The aerodynamic restoring moments tend to impart a restoring spring characteristic with respect to the craft pitch attitude.
It is known in the art to enhance the oscillatory dynamic stability of the craft by utilizing stability augmentation systems. Such systems generally utilize pitch rate in the control laws. The sensors commonly utilized to measure this parameter are angular accelerometers or rate gyros. Present day angular accelerometers suitable for control purposes tend to be expensive and generally the signals provided thereby are only suitable for oscillatory high frequency damping. Present day rate gyroscopes as well as tending to be expensive, have relatively low long term reliability because of the rapidly rotating members.
Such stability augmentation systems utilized for oscillatory damping purposes have heretofore not been flight critical components since even if total failure of the system should occur the aircraft would retain its aerodynamic longitudinal static stability which imparted safe flight and adequate pilot handling characteristics to the aircraft so that the flight could be safely completed.
It is appreciated that in order to obtain the relatively large balancing moments about the craft pitch axis to provide longitudinal static stability, considerable drag is introduced that adversely affects the fuel economy of the aircraft. It is currently being considered to relax the aerodynamic longitudinal static stability of the aircraft so as to increase fuel economy. This would be accomplished by maintaining the horizontal stabilizer of the aircraft in a relatively aerodynamically streamlined position and perhaps by reducing the surface area thereof. It would be necessary then to design the aircraft with the center of lift close to the center of gravity. Since the craft would no longer have the stiff spring aerodynamic restoring moments that heretofore imparted static stability to the craft, the marginally stable or statically unstable craft would no longer have safe flight characteristics and desirable pilot handling characteristics. In such an aircraft a gust disturbance causing a pitch deflection may result in the craft continuing to diverge in pitch attitude. The pilot manual controls of such a craft may be extremely sensitive whereby it the pilot should impart a pitch control motion to the column, the craft may respond with an excessive pitch maneuver leading to a tendency toward pilot induced oscillations.
In order to render safe the flight characteristics of a relaxed longitudinal static stability aircraft and to provide desirable pilot handling characteristics, it is necessary to replace the aerodynamic static stability with a stability augmentation system. In the instance of a relaxed static stability aircraft, however, the stability augmentation system becomes a flight critical component of the aircraft whereby total failure thereof could result in loss of the craft. In order to impart the necessary reliability to such systems, generally triply or quadruply redundant channels are required. In order to provide the necessary static stability for the relaxed static stability aircraft, a control law that includes a pitch attitude term as well as terms measuring pitch rate, pitch acceleration and vertical acceleration are required. Since attitude sensors such as vertical gyroscopes or inertial platforms tend to be expensive, heavy and subject to failure and rate gyroscopes and angular accelerometers have the disadvantages discussed above, quadruply redundant sensor instrumentation for such a system would tend to be prohibitively expensive, bulky and heavy while tending to be unreliable because of the rapidly rotating members utilized in such gyroscopic instrumentation.
Linear accelerometers on the other hand are relatively inexpensive and highly reliable since such instruments do not include rapidly rotating components. As well as measuring linear vertical acceleration, a pair of linear accelerometers mounted respectively forward and aft of the center of gravity or merely separated by a reasonable distance have been utilized to provide a measure of pitch acceleration and pitch rate. Such use of linear accelerometers is taught in Applicant's assignee's U.S. Pat. Nos. 3,007,656, entitled "Aircraft Automatic Pilot" by H. Miller et al, issued Nov. 7, 1961; 2,808,999 entitled "Automatic Flight Control Apparatus" by P. J. Chenery, issued Oct. 8, 1957; and 2,487,793 entitled "Object Controlling Electric Motor System", by O. E. Esval et al, issued Nov. 15, 1949. In such prior art instrumentation, e.g., U.S. Pat. No. 3,007,656, the outputs of the linear accelerometers are combined to provide pitch angular acceleration, which combined signal is passed through a lag network to provide a signal that simulates pitch angular rate with a washout.
In the prior art linear accelerometers have only been utilized for dynamic high frequency stabilization. This was the situation because the linear accelerometers separately, and combined as described in U.S. Pat. No. 3,007,656, are subject to gradient errors and bias errors which are necessarily washed out in using these sensors. Thus prior to the present invention linear accelerometers with their steady state errors could not be utilized to provide an accurate and reliable measure of attitude which measure is required for static stabilization in a stability augmentation system for relaxed static stability aircraft.
It is the desideratum of the present invention to provide a suitable stability augmentation system for relaxed static stability aircraft utilizing linear accelerometers as the primary sensing instruments.